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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 235 records · Page 13
All sourcesNASA NTRSDOE OSTI
NASA NTRS1956

Effect of design over-all compressor pressure ratio division on acceleration characteristics of three hypothetical two-spool turbojet engines

The engines, each with a compressor overall total-pressure ratio of 12 and a design inner-turbine-inlet temperature of 2500 degrees R, were investigated at static sea-level conditions to determine the effect on transient performance of varying the desitn pressure ratio divisions 2-6, 3-4, and 4-3 between the outer and inner compressors. The transient considered was an acceleration from 40 to 100 percent design thrust. When the outer compressor of each engine reached design speed, the inner compressors were overspeeding, the maximum being only 1.7 over design mechanical speed. Acceleration times for the three engines were equal.

TURBINES - MATCHINE↗
NASA NTRS1996

Three-Dimensional Flow Field Measurements in a Transonic Turbine Cascade

Three-dimensional flow field measurements are presented for a large scale transonic turbine blade cascade. Flow field total pressures and pitch and yaw flow angles were measured at an inlet Reynolds number of 1.0 x 10(exp 6) and at an isentropic exit Mach number of 1.3 in a low turbulence environment. Flow field data was obtained on five pitchwise/spanwise measurement planes, two upstream and three downstream of the cascade, each covering three blade pitches. Three-hole boundary layer probes and five-hole pitch/yaw probes were used to obtain data at over 1200 locations in each of the measurement planes. Blade and endwall static pressures were also measured at an inlet Reynolds number of 0.5 x 10(exp 6) and at an isentropic exit Mach number of 1.0. Tests were conducted in a linear cascade at the NASA Lewis Transonic Turbine Blade Cascade Facility. The test article was a turbine rotor with 136 deg of turning and an axial chord of 12.7 cm. The flow field in the cascade is highly three-dimensional as a result of thick boundary layers at the test section inlet and because of the high degree of flow turning. The large scale allowed for very detailed measurements of both flow field and surface phenomena. The intent of the work is to provide benchmark quality data for CFD code and model verification.

Giel, P. W.↗
NASA NTRS1995

A Numerical Study of the Effect of Wake Passing on Turbine Blade Film Cooling

Time-accurate and steady three-dimensional viscous turbulent numerical simulations were performed to study the effect of upstream blade wake passing unsteadiness on the performance of film cooling on a downstream axial turbine blade. The simulations modeled the blade as spanwise periodic and of infinite span. Both aerodynamic and heat transfer quantities were explored. A showerhead film cooling arrangement typical of modern gas turbine engines was employed. Showerhead cooling was studied because of its anticipated strong sensitivity to upstream flow fluctuations. The wake was modeled as a region of zero axial velocity on the upstream computational boundary which translated with each iteration. This model is compatible with a planned companion experiment in which the wakes will be produced by a rotating row of cylindrical rods upstream of an annular turbine cascade. It was determined that a steady solution with appropriate upstream swirl and stagnation pressure predicted the span-average film effectiveness quite well. The major difference is a 2 to 3 percent overprediction of span-average film effectiveness by the steady simulation on the pressure surface and in the showerhead region. Local overpredictions of up to 8 percent were observed in the showerhead region. These differences can be explained by the periodic relative lifting of the boundary layer and enhanced mixing in the unsteady simulations.

Heidmann, James D.↗
NASA NTRS1992

Investigation of rotor blade roughness effects on turbine performance

The cold air test program was completed on the SSME (Space Shuttle Main Engine) HPFTP (High Pressure Fuel Turbopump) turbine with production nozzle vane rings and polished coated rotor blades with a smooth surface finish of 30 microinch (0.76 micrometer) RMS (Root Mean Square). The smooth blades were polished by an abrasive flow machining process. The test results were compared with the air test results from production rough coated rotor blades with a surface finish of up to 400 microinch (10.16 micrometer) RMS. Turbine efficiency was higher for the smooth blades over the entire range tested. Efficiency increased 2.1 percentage points at the SSME 104 percent RPL (Rated Power Level) condition. This efficiency improvement could reduce the SSME HPFTP turbine inlet temperature by 57 degrees Rankine (32 degrees Kelvin) increasing turbine durability. The turbine flow parameter increased and the mid-span outlet swirl angle became more axial with the smooth rotor blades.

Boynton, J. L.↗
NASA NTRS1987

The determination of equivalent bearing loading for the BSMT that simulate SSME high pressure oxidizer turbopump conditions using the SHABERTH/SINDA computer programs

The MSFC bearing seal material tester (BSMT) can be used to evaluate the SSME high pressure oxygen turbopump (HPOTP) bearing performance. The four HPOTP bearings have both an imposed radial and axial load. These radial and axial loads are caused by the HPOTP's shaft, main impeller, preburner impeller, turbine and by the LOX coolant flow through the bearings, respectively. These loads coupled with bearing geometry and operating speed can define bearing contact angle, contact Hertz stress, and heat generation rates. The BSMT has the capability of operating at HPOTP shaft speeds, provide proper coolant flowrates but can only apply an axial load. Due to the inability to operate the bearings in the BSMT with an applied radial load, it is important to develop an equivalency between the applied axial loads and the actual HPOTP loadings. A shaft-bearing-thermal computer code (SHABERTH/SINDA) is used to simulate the BSMT bearing-shaft geometry and thermal-fluid operating conditions.

Mcdonald, Gary H.↗
NASA NTRS1983

A comprehensive method for preliminary design optimization of axial gas turbine stages. II - Code verification

The present effort represents an extension of previous work wherein a calculation model for performing rapid pitchline optimization of axial gas turbine geometry, including blade profiles, is developed. The model requires no specification of geometric constraints. Output includes aerodynamic performance (adiabatic efficiency), hub-tip flow-path geometry, blade chords, and estimates of blade shape. Presented herein is a verification of the aerodynamic performance portion of the model, whereby detailed turbine test-rig data, including rig geometry, is input to the model to determine whether tested performance can be predicted. An array of seven (7) NASA single-stage axial gas turbine configurations is investigated, ranging in size from 0.6 kg/s to 63.8 kg/s mass flow and in specific work output from 153 J/g to 558 J/g at design (hot) conditions; stage loading factor ranges from 1.15 to 4.66.

Jenkins, R. M.↗
NASA NTRS1993

Two stage turbine for rockets

The aerodynamic design and rig test evaluation of a small counter-rotating turbine system is described. The advanced turbine airfoils were designed and tested by Pratt & Whitney. The technology represented by this turbine is being developed for a turbopump to be used in an advanced upper stage rocket engine. The advanced engine will use a hydrogen expander cycle and achieve high performance through efficient combustion of hydrogen/oxygen propellants, high combustion pressure, and high area ratio exhaust nozzle expansion. Engine performance goals require that the turbopump drive turbines achieve high efficiency at low gas flow rates. The low mass flow rates and high operating pressures result in very small airfoil heights and diameters. The high efficiency and small size requirements present a challenging turbine design problem. The shrouded axial turbine blades are 50 percent reaction with a maximum thickness to chord ratio near 1. At 6 deg from the tangential direction, the nozzle and blade exit flow angles are well below the traditional design minimum limits. The blade turning angle of 160 deg also exceeds the maximum limits used in traditional turbine designs.

Veres, Joseph P.↗
NASA NTRS1992

Investigation of three-dimensional flow field in a turbine including rotor/stator interaction. II - Three-dimensional flow field at the exit of the nozzle

The nozzle exit flowfield was measured at two axial locations with a miniature five-hole probe. Measurements were taken from hub-to-tip, blade-to-blade at 21 radial locations and at two axial locations downstream of the nozzle trailing edge to resolve the flowfield accurately including the nozzle wake, secondary flow region, horseshoe vortex and losses. All three components of the velocity, stagnation pressure, static pressure, and pitch and yaw angles have been resolved very accurately. The wake data seems to indicate that the decay of the wake is faster than the wake of an isolated nozzle row. The cause of this is attributed to the presence of the rotor downstream. A distinct vortex core has been observed near the tip. The indications are that the horseshoe vortex and the passage vortex have merged to produce a single loss core region. Roughly a third of the blade height passage near the tip and a third of the blade height near the hub is dominated by secondary flow, passage vortex and the horseshoe vortex phenomena. Only the middle third of the nozzle behaves as per design. These and other data are presented, interpreted and synthesized to understand the nozzle flowfield.

Zaccaria, M.↗
NASA NTRS1954

Aspects of internal-flow-system design for helicopter propulsive units

Pertinent items relating to the design of internal-flow systems for reciprocating engine, turbine engine, and pressure-jet installations in helicopters are discussed. The importance of the following items is emphasized: controllable exit and ram recovery for reciprocating-engine cooling; performance penalties possible in turbine-engine installations attributed to distortions in the flow distribution at the face of the axial compressor; the effects of high subsonic flow velocities on the performance of duct elements; and the effects of centrifugal forces on the flow in a ducted helicopter rotor. Data on the maximum flow capacities of bends are included.

John R Henry↗
DOE OSTI2023

Development of a Large-Scale Open Geometry for Squealer Tip Coolant Injection Research

In this work, a large-scale symmetric canonical axial turbine squealer tip model is developed, which allows for the investigation of incoming gas path flow interaction with coolant flow injected into the squealer cavity. The model is developed to investigate Reynolds numbers on the order of one million for their applicability to novel supercritical CO2 cycles.

Tuite, Logan M.↗
NASA NTRS1977

Pollution reduction technology program for class T4(JT8D) engines

The technology required to develop commercial gas turbine engines with reduced exhaust emissions was demonstrated. Can-annular combustor systems for the JT8D engine family (EPA class T4) were investigated. The JT8D turbofan engine is an axial-flow, dual-spool, moderate-bypass-ratio design. It has a two-stage fan, a four-stage low-pressure compressor driven by a three-stage low-pressure turbine, and a seven-stage high-pressure compressor driven by a single-stage high-pressure turbine. A cross section of the JT8D-17 showing the mechanical configuration is given. Key specifications for this engine are listed.

Roberts, R.↗
NASA NTRS2010

Controlling Separation in Turbomachines

Four examples of flow control: 1) Passive control of LP turbine blades (Laminar separation control). 2) Aspiration of a conventional axial compressor blade (Turbulent separation control). 3) Compressor blade designed for aspiration (Turbulent separation control). 4.Control of intakes in crosswinds (Turbulent separation control).

Evans, Simon↗
NASA NTRS2012

Comparison of Various Supersonic Turbine Tip Designs to Minimize Aerodynamic Loss and Tip Heating

The rotor tips of axial turbines experience high heat flux and are the cause of aerodynamic losses due to tip clearance flows, and in the case of supersonic tips, shocks. As stage loadings increase, the flow in the tip gap approaches and exceeds sonic conditions. This introduces effects such as shock-boundary layer interactions and choked flow that are not observed for subsonic tip flows that have been studied extensively in literature. This work simulates the tip clearance flow for a flat tip, a diverging tip gap and several contoured tips to assess the possibility of minimizing tip heat flux while maintaining a constant massflow from the pressure side to the suction side of the rotor, through the tip clearance. The Computational Fluid Dynamics (CFD) code GlennHT was used for the simulations. Due to the strong favorable pressure gradients the simulations assumed laminar conditions in the tip gap. The nominal tip gap width to height ratio for this study is 6.0. The Reynolds number of the flow is 2.4 x 10(exp 5) based on nominal tip width and exit velocity. A wavy wall design was found to reduce heat flux by 5 percent but suffered from an additional 6 percent in aerodynamic loss coefficient. Conventional tip recesses are found to perform far worse than a flat tip due to severe shock heating. Overall, the baseline flat tip was the second best performer. A diverging converging tip gap with a hole was found to be the best choice. Average tip heat flux was reduced by 37 percent and aerodynamic losses were cut by over 6 percent.

Shyam, Vikram↗
NASA NTRS1983

Axial vane-type swirler performance characteristics

The performance of an axial vane-type swirler was investigated to aid in computer modeling of gas turbine combustor flowfields and in evaluation of turbulence models for swirling confined jet flow. The swirler studied is annular with a hub-to-swirler diameter ratio of 0.25 and ten adjustable vanes of pitch-to-chord ratio 0.68. Measurements of time-mean axial, radial, and tangential velocities were made at the swirler exit plane using a five-hole pitot probe technique with computer data reduction. Nondimensionalized velocities from both radial and azimuthal traverses are tabulated and plotted for a range of swirl vane angles phi from 0 to 70 degrees. A study was done of idealized exit-plane velocity profiles relating the swirl numbers S and S' to the ratio of maximum swirl and axial velocities for each idealized case, and comparing the idealized swirl numbers with ones calculated from measured profiles.

Sander, G. F.↗
NASA NTRS1996

The Effect of Vane-Blade Spacing on Transonic Turbine Stage Performance

This paper presents the results of a computational study on the effect of axial spacing between the vane and blade rows of a transonic turbine stage. The study was performed on the mid-span section of a high-pressure turbine stage using a quasi-3D, unsteady Navier-Stokes solver that provides a fully interactive vane-blade unsteady flow solution. Three different cases were considered, corresponding to axial spacings of 20%, 40%, and 60% of the vane axial chord. The calculated vane and blade pressure distributions for the 40 percent case were found to compare favorably with experimental measurements acquired in a short-duration shock tunnel. In addition, the analysis shows a marked increase in the amplitude of the unsteady pressure fluctuations on the vane and blade surfaces as the spacing decreases. Time-averaged stage adiabatic efficiency predictions for each case are presented to show the effect of spacing on aerodynamic performance.

F J Kelecy↗
NASA NTRS1988

The simulation of the alternate turbopump development high pressure oxygen and fuel turbopumps for the space shuttle main engine using the Shaberth computer program

The Space Shuttle Main Engine (SSME) is basically comprised of a combustion chamber and nozzle, high and low pressure oxygen turbopumps and high and low pressure fuel turbopumps. In the current configuration, the high pressure fuel (HPTFP) and high pressure oxygen turbopumps (HPOTP) have experienced a history of ball bearing wear. The wear problem can be attributed to numerous factors including the hydrodynamic axial and radial loads caused by the flow of liquid oxygen and liquid hydrogen through the turbopump impellers and turbine. Also, friction effects between the rolling elements, races, and cage can create thermally induced bearing geometry changes. To alleviate some of the current configuration problems, an alternate turbopump development (ATD) was proposed. However, the ATD HPOTP and HPTFP are constrained to operate interchangeably with the current turbopumps, thus, the operation conditions must be similar. The ATD configuration features a major change in bearings used to support the integrated shaft, impeller, and turbine system. A single ball and single roller will replace the pump-end and turbine and duplex ball bearings. The Shaft-Bearing-Thermal (SHABERTH) computer code was used to model the ATD HPOTP and ATD HPFTP configurations. A two bearing model was used to simulate the HPOTP and HPFTP bearings and shaft geometry. From SHABERTH, a comparison of bearing reaction loads, frictional heat generation rates, and Hertz contact stresses will be attempted with analysis at the 109 percent and 65 percent power levels.

Mcdonald, Gary H.↗
NASA NTRS1989

Critical Evaluation of the Conference on Low Reynolds Number Aerodynamics at the University of Notre Dame

As a result of the continued interest in designing efficient low Reynolds number systems, the University of Notre Dame decided to organize a Conference on Low Reynolds Number Aerodynamics in June 1989. This Conference followed the 1986 International Conference in London by about three years and the first Conference on Low Reynolds Number Airfoil Aerodynamics at Notre Dame in 1985 by four years. The emphasis of the 1989 Conference was to assess the state-of-the-art in the chord Reynolds number range from about 10,000 to about 700,000. Applications of current interest include high altitude remotely or robotically piloted vehicles, ultra-light, and human powered vehicles as well as mini-RPVs at low altitudes. Other examples include small axial-flow fans used to cool electronic equipment in the unpressurized sections of high-altitude aircraft and gas turbine blades. High Reynolds number airfoil design strategies attempt to control the onset and development of turbulent boundary layers. This is difficult at low Reynolds numbers because of the increased stability of attached laminar boundary layers. Therefore, laminar separation is common even at small angles of attack at low Reynolds numbers. Under these conditions, the development of a turbulent boundary layer usually depends on the formation of a transitional separation bubble. The purpose of this Conference on Low Reynolds Number Aerodynamics was to bring together those researchers who have been active in areas closely related to this subject. It is clear from the papers presented that a great deal of progress has been made in understanding the occurrence and behavior of laminar separation and transition as well as their overall effect on the performance of airfoils at low chord Reynolds numbers. This progress has brought us closer to our goal of improving analytical methods for the design and evaluation of a variety of practical applications.

Mueller, Thomas J.↗
NASA NTRS1981

Liquid oxygen turbopump technology

A small, high-pressure, LOX turbopump was designed, fabricated and tested. The pump is a single-stage centrifugal type with power to the pump supplied by a single-stage partial-admission axial-impulse turbine. Design conditions included an operating speed of 7330 rad/s (70,000 rpm), pump discharge pressure of 2977 N/sqcm (4318 psia), and a pump flowrate of 16.4 Kg/s (36.21 lb/s). The turbopump contains a self-compensating axial thrust balance piston to eliminate axial thrust loads on the bearings during steady-state operation. Testing of the turbopump was achieved usng a gaseous hydrogen high-pressure flow to drive the turbine, which generally is propelled by LOX/LH2 combustion products, at 1041K (1874 R) inlet temperature and at a design pressure ratio of 1.424. Test data obtained with the turbopump are presented which include head-flow-efficiency performance, suction performance, balance piston performance and LOX seal performance. Mechanical performance of the turbopump is also discussed.

Nielson, C. E.↗